Gas turbines, which (for example) are used in terrestrial power generation and jet/propjet aircraft propulsion, may burn fuel having an adiabatic flame temperature greater than the melting or softening point of parts such as turbine blades, turbine inlet guide vanes, turbine stators, turbine shaft, combustor walls, and/or other parts to which hot combustion products are exposed; and which may be individually or collectively referred to as (a) high temperature surface(s). Various approaches have been used to address this apparent barrier. For example, development of high temperature alloys has historically represented significant research emphasis. Air cooling has been used in various forms to provide convective heat transfer from heated surfaces.
While air cooling of some gas turbine parts has resulted in increases in allowable combustion gas temperature, achievable increases in temperature have been limited and generally do not allow for operation at full adiabatic flame temperature. For example, surface cooling has been less than optimal because cooling effects have been limited to small radii surrounding surface-penetrating fluid passages. For example, in turbine blade cooling applications, cooling air exits surface-penetrating cooling holes at relatively high velocity. While providing convective cooling to the inner diameter of the cooling holes themselves, the high velocity air also tends to disturb flow over the surface of the turbine blade, ultimately causing vortices that can bring hot combustion gases into contact with turbine blade surfaces peripheral to the cooling holes.
Combustion gas dilution and/or other stoichiometric imbalances have been used to reduce the temperature of combustion gases. Unfortunately, gas turbine thermodynamic efficiency is negatively affected by dilution the heat source with cooling air, because cooling air lowers the peak cycle temperature of the heat engine.
Thermodynamic efficiency may be expressed as a function of the ratio of heat source temperature to heatsink temperature. The heatsink temperature may be very close to or substantially equal to ambient air temperature (effective heat sink temperature may be influenced by flow losses). The heat source temperature has typically been limited, not by adiabatic flame temperature, but by material property limitations (as described above). Dilution of the combustion gases, and corresponding decrease in heat source temperature may cause some aircraft gas turbines such as turbofan engines to be limited in thermodynamic efficiency to around 37% maximum. Terrestrial gas turbines, such as those used to provide peak power to electric power grids, may typically reach a thermodynamic efficiency of around 60%, but must use a topping cycle to remove energy from the combustion gases in stages to limit peak temperatures. Topping cycle equipment (which may, for example, use non-stoichiometric, staged combustion) is generally associated with high capital costs and relatively large size.
Notwithstanding incorrect prior art references to “film-cooling”, attempts to maintain an actual cooling film adjacent to high temperature surfaces has apparently not been tried, because cooling flow has been in the form of jets that, as described above, actually increase vorticity and heat flow to portions of heated surfaces peripheral to cooling holes.
What is needed is technology that allows a gas turbine to operate with higher thermodynamic efficiency by reducing dilution and/or topping cycles, while protecting high temperature surfaces such as turbine blades, combustor walls, turbine inlet guide vanes, turbine shaft, turbine stators and/or other heat-exposed parts from the increased temperature. What is also needed is a technology that may reduce gas turbine capital cost.